Quest for Performance: The Evolution of Modern Aircraft
Chapter 11: Early Jet Fighters
Contemporary Fighters
[328] The decade of the 1970's saw the introduction of four new fighters into the armed services of the United States. Entering the Navy was the Grumman F-14A Tomcat, while the McDonnell F-15 Eagle and the General Dynamics F-16 Fighting Falcon became part of the USAF inventory. Finally, the unique British Aerospace AV-8A Harrier vertical takeoff and landing (VTOL) fighter entered service with the U.S. Marine Corps. These aircraft are described below. Since these machines are currently first-line equipment in the United States inventory, only limited performance data are available for them in the open literature. Hence, performance and aerodynamic data in table V for these aircraft are limited.
Grumman F-14A Tomcat
In January 1969 the Grumman Aerospace Corporation was named the winner in a design competition for development of a Navy fighter to fill the role for which the F-111 was rejected. First flight of the new fighter, known as the F-14A Tomcat, took place on December 21, 1970, and the first operational squadrons were deployed on the U.S.S. Enterprise in September 1974. In addition to the previously described combat-air-patrol (CAP) mission, the F-14A was designed to fill several other roles including escort of carrier-launched strike forces, deck-launched interceptions, close-in air-to-air combat, and low-altitude strike missions. These varied missions spelled the need for an aircraft [329] with a combination of high cruise efficiency at subsonic speeds, good maneuverability at high-subsonic/transonic speeds, and a supersonic capability extending to Mach 2.4. Finally, as in all Navy fighter aircraft, low approach speeds compatible with carrier operations were required. As a consequence, it came as no surprise that the F-14A turned out to be an aircraft featuring variable-sweep wings.
Photographs of the F-14A are presented in figures 11.30 and 11.31, and the wing-planform shape is shown in figure 11.32; physical and performance data for the aircraft are contained in table V. As compared with the variable-sweep F-111, the Tomcat has distinct differences in appearance. Among the distinguishing features of the F-14A are the large two-dimensional horizontal-ramp supersonic inlets. In accordance with the Mach number, the angle of the upper ramp, that is, the inside horizontal surface of the upper part of the inlet, varies automatically at supersonic speeds to maintain high inlet pressure recovery. Another identifying feature of the aircraft is the two vertical-tail units necessary for adequate directional stability and control at high angles of attack and high Mach numbers. The crew of the Tomcat is accommodated in a tandem arrangement, in contrast to the side-by-side seating in the F-111.
An examination of the physical data in table V shows that the F-14A is significantly lighter than the F-111 and has a lower wing loading, a higher thrust-to-weight ratio, and a much shorter length. All....

aerial view of a F-14A
Figure 11.30 - In-flight view of Grumman F-14A Tomcat twin-engine jet fighter. [ukn] [Original photo was in color, Chris Gamble, html editor]


ground view of a F-14A being serviced by ground crew
[330] Figure 11.31 - Grumman F-14A Tomcat twin-engine jet fighter. [author's collection]


....these differences increase carrier compatibility. Two Pratt & Whitney TF30-P-412A afterburning turbofan engines power the F-14A; this is a version of the same engine used in the F-111. Repowering the aircraft with a more modern engine was originally planned; but so far, this has not taken place.

Wing sweepback angle of the Tomcat varies in flight from 20° to 68° to decrease the space required for storage on the aircraft carrier, the wing span is further reduced by increasing the wing sweepback angle to 75°. Wing thickness ratio (in the streamwise direction) varies from 9 percent for the low sweep position to 5 percent for a sweep angle of 68. An important difference in the wing geometry of the F-14 and F-111 is shown in figures 11.28 (F-111A) and 11.32 (F-14A). In terms of the wing semispan in the low sweep position, the pivot of the F- 14A is 10 to 12 percent farther outboard than that of the F-111. According to the paper by Kress in reference 155, the more outboard pivot location results in a much reduced rearward movement of the center of lift with increasing sweep angle. As a consequence, trim drag is reduced and available pitch-control power is increased. The favorable effect of locating the pivot in the proper outboard position is, of course, in accordance with NASA basic research. (See chapter 10.) An interesting feature of the F-14A wing is the retractable vane located on the fixed portion of the wing; the vane is shown in figure 11.32 in both [331] the retracted (low wing-sweep) and extended (high wing-sweep) positions. The function of the vane is to reduce the rearward shift in the center of lift that accompanies an increase in Mach number from subsonic to supersonic values. (See figure 10.15.)
Leading-edge slats and trailing-edge flaps are used to improve maneuverability at high subsonic speeds as well as to increase wing maximum lift coefficient at low speeds. The auxiliary flap shown in figure 11.32 is used only at low speeds to increase maximum lift. In normal operation, the maneuvering flaps, wing-sweep angle, and vane position are automatically controlled by a computer in accordance with a stored program that utilizes inputs from several measured flight parameters such as angle of attack, static and total pressures, and temperature. Manual operation of the wing is also possible. Roll control of the aircraft is provided by a combination of wing spoilers and differential deflection of the horizontal-tail surfaces.
Although available performance information on the F-14A is sketchy, the data in table V show maximum Mach numbers of 2.4 at...

ilustration of F-14A wing sweep area
Figure11.32 - Approximate wing-planform shape of Grumman F-14A variable-sweep jet fighter.


[332] 49 000 feet and 1.2 at sea level and a time of only 2.1 minutes required to reach an altitude of 60 000 feet. According to the Kress paper in reference 155, the maximum subsonic lift-drag ratio is about 15, which is much higher than the value of 8.58 given in table V for the F-4.

The F-14 is armed with the Vulcan 20-mm rotary cannon for close-in combat and, depending on the mission, can carry a combination of Sidewinder, Sparrow, and Phoenix missiles. As many as six of the latter missiles can be carried on a combat-air-patrol mission. The attack radar is capable of tracking simultaneously 24 separate targets at ranges as great as 100 miles. (Drones have been hit at ranges of over 100 miles with the Phoenix missile in practice missions.) All six of the Phoenix missiles can be fired together, and each can be guided to a different target. With a range of only 2 miles, the Sidewinder is used in short-range air-to-air combat, while the Sparrow with a 10-mile range is employed for more distant engagements. All the missiles are carried externally, but none are attached to the movable portions of the wing; hence, the complication of swiveling store-mounting pylons, such as used on the F-111, is avoided.
The Tomcat appears to be a fighter with very high performance and great operational versatility. By the beginning of 1980, about 400 F-14 aircraft had been built. Included in this group were 80 units for Iran. The aircraft is still in production and is likely to remain so for a number of years.
McDonnell Douglas F-15 Eagle
Experience in the Vietnam conflict showed the F-4 Phantom II to have maneuvering performance inferior to that of the Soviet-built MiG21. In response to this finding, the USAF developed a set of requirements for a dedicated air-superiority fighter with a maneuvering capability greater than any existing or foreseeable-future fighter aircraft. McDonnell Douglas, North American Rockwell, and Fairchild-Republic submitted proposals in the ensuing design competition. McDonnell Douglas was chosen as the winner in late 1969, and the F-15 Eagle made its first flight on July 27, 1972. By mid-1980, 941 of these aircraft had been built or were on order, including units for Israel, Japan, and Saudi Arabia as well as those for the USAF. In addition to United States production, the aircraft is also being manufactured under license in Japan.
[333] To understand the design of the F-15 and its unique capabilities, some insight into the meaning of maneuverability and its relation to several aircraft design parameters is necessary.
Aircraft Maneuverability
The maneuvering capability of an aircraft has many facets, but one of the most important of these is its turning capability. In a combat situation between two opposing fighters flying at the same speed, the aircraft capable of turning with the shortest radius of turn without losing altitude usually has the advantage. This assumes equality of many other factors such as aircraft stability and control characteristics, armament, and, of course, pilot skill.
In steady, turning flight the lift developed by the wing must balance not only the weight of the aircraft but the centrifugal force generated by the turn. (The term "balance" is used here in a vector sense; that is, the lift vector must equal the sum of the weight and centrifugal force vectors.) The load factor is defined as the ratio of the lift in the turn to the weight of the aircraft and is usually expressed in g units, where g is the acceleration due to gravity. Thus, a 2-g turn is one in which the wing must develop a lift force twice the weight of the aircraft. The value of the load factor is uniquely defined by the aircraft angle of bank. For example, 2-g and 5-g turns require bank angles of 60 and 78.5 respectively. Finally, for a given bank angle and thus load factor, the turning radius varies as the square of the speed; for example, doubling the speed of the aircraft increases the turning radius by a factor of 4. It would then appear that two different aircraft flying at the same speed would have the same turning radius; however, this conclusion is not necessarily correct. The maximum load factor and associated turning radius may be limited by wing stalling. For a given speed and altitude, stalling occurs as a function of the wing maximum lift coefficient and the wing loading in straight and level flight. Clearly then, the turning capability of different aircraft types may vary widely.
To give some physical significance to these qualitative ideas on turning performance, the variation with altitude of the maximum achievable load factor is shown in figure 11.33(a) for wing loadings of 50 and 100 pounds per square foot; the corresponding variation of turning radius with altitude is shown in figure 11.33(b) for the same two wing loadings. The curves were calculated for a Mach number of 0.85 and a limit maximum lift coefficient of 0.70. For a wing loading of....

graph of altitude vs load  at 50 & 100 pouds per square foot
(a) Variation of normal load factor with wing loading and altitude.
graph of turningradius vs 50 psf and 100  psf
(b) Variation of turning radius with wing loading and altitude.
[334] Figure 11.33 - Effect of wing loading and altitude on turning performance of a fighter aircraft. M = 0.85, CL = 0.7, constant altitude.

[335] ....50 pounds per square foot, the structural design load limit of 7.33 g can be realized up to an altitude of nearly 15 000 feet; after which, the maximum lift capability of the wing limits the allowable load factor until only 2.3 g can be achieved at 40 000 feet. In comparison, the maximum lift capability limits the achievable load factor of the aircraft with a wing loading of 100 pounds per square foot at all altitudes, with straight and level flight at 1g being just barely possible at 40 000 feet. The corresponding effects of wing loading and altitude on turning radius are equally dramatic, as shown in figure 11.33(b).
The effect of wing maximum lift coefficient is inverse to that of wing loading; that is, increasing the maximum lift coefficient acts in the same way as reducing the wing loading. For example, increasing the maximum lift coefficient from 0.7 to 1.4 would shift the curve for a wing loading of 100 pounds per square foot to the exact position as that occupied by the curve for a wing loading of 50 pounds per square foot and a maximum lift coefficient of 0.7. At a given Mach number, the maximum lift coefficient depends upon the wing-planform shape, the airfoil section, and the type of maneuvering flaps used if any. To further complicate the picture, the maximum lift coefficient also varies with Mach number in a manner that again depends upon the wing design parameters. The message of figure 11.33, however, is quite clear: the turning performance improves as the wind loading decreases and the maximum lift coefficient increases.
Two other important aircraft physical parameters may also limit turning performance. First, at a given speed and altitude, the aircraft drag increases rapidly with lift coefficient; as a consequence, the available thrust may not be sufficient to balance the drag at some load factors that the wing can sustain. In this case the aircraft loses altitude in the turn, an undesirable situation in combat. As for maximum lift coefficient, the drag rise with increasing lift depends upon the wing design and Mach number, as well as upon the added drag required to trim the aircraft at high lift coefficients. Finally, the turning performance may be limited by the control power available in the horizontal tail for trimming the aircraft at the high maneuvering lift coefficients.
These ideas are embodied in a technique for describing and specifying fighter aircraft maneuverability. Known by the term "energy maneuverability," the technique involves the specification of desired aircraft climb and/or acceleration capability for various combinations of speed, altitude, and turning load factor. The quantity specified for each of these combinations is labeled "specific excess power" Ps and is simply the excess power available per unit aircraft weight as compared [336] with the power required to maintain constant altitude in the turn. As an illustration, a value of Ps of 500 might be specified for a 6-g turn at Mach 0.6 at an altitude of 25 000 feet. This simply means that sufficient power is available in the aircraft to establish a steady rate of climb of 500 feet per minute while maintaining the specified turn. Alternatively, the excess power could be used to accelerate to a higher speed while in the turn.
For the first time, extensive use was made of the energy-maneuverability technique in establishing the requirements that led to the McDonnell Douglas F-15 Eagle air-superiority fighter. Values of P were specified for 13 combinations of speed, altitude, and g-loading at subsonic, transonic, and supersonic Mach numbers for the new fighter. In addition, maximum and cruise Mach numbers were specified, as well as landing, takeoff, and range requirements.
Aircraft Description
The McDonnell Douglas F-15 Eagle emerged from the complex and extensive set of requirements established by the USAF. Views of the all-weather single-place fighter are shown in figures 11.34 and 11.35, and a sketch of the wing-planform shape is given in figure 11.36. Configuration of the twin-engine aircraft is characterized by a high-mounted wing, twin vertical tails mounted at the rear of the short fuselage, and large, horizontal-ramp variable- geometry external-compression inlets located on the sides of the fuselage ahead of the wing. The horizontal-tall surfaces are mounted in the low position on fuselage extensions on either side of the exhaust nozzles.
The data in table V show that, based on the design maximum weight, the wing loading of the F-15 is significantly lower and the thrust loading much greater than corresponding values for earlier fighter aircraft. At the lower weights to be expected during combat, wing loadings as low as 55 pounds per square foot and static thrust-to-weight-ratios of as much as 1.35 might be expected. (As the Mach number increases at a given altitude, the thrust of the afterburning turbofan also increases. For example, the thrust of the F-15 engine at sea level and Mach 0.9 is nearly twice the sea-level static value.) The values of these parameters represent a significant departure from previous fighter design philosophy and resulted from the energy-maneuverability concepts employed in specifying the aircraft. Note that even at design takeoff weight, the aircraft is capable of sustained vertical flight. Maximum speeds are listed as Mach 2.54 and 1.21 at 40 000 feet and....

aerial view of F-15
[337] Figure 11.34 - McDonnell Douglas F-15 Eagle twin-engine jet fighter. [mfr] [Original photo was in color, Chris Gamble, html editor]


....sea level, respectively. Service ceiling is 63 000 feet, and ferry range with maximum external fuel is 3570 miles. No other performance information is available, but the aircraft undoubtedly has outstanding performance and maneuvering capability.

The wing planform of the F-15, shown in figure 11.36, suggests a modified cropped delta shape with a leading-edge sweepback angle of 45°. Ailerons and a simple high-lift flap are located on the trailing edge. No leading-edge maneuvering flaps are utilized, although such flaps were extensively analyzed in the design of the wing. This complication was avoided, however, by the combination of low wing loading and fixed leading-edge camber that varies with spanwise position along the wing. Airfoil thickness ratios vary from 6 percent at the root to 3 percent at the tip. An interesting discussion of the wing design and the many trade-off studies involved in its finalization are contained in a paper by Niedling included in reference 155.
Propulsion of the F-15 is supplied by two Pratt & Whitney F100-PW-100 afterburning turbofan engines of 23 904/14 780 pounds thrust each. Developed especially for the F-15, these high-pressure-ratio engines are reported to have much improved efficiency over earlier engines for fighter aircraft.

overhead aerial view of F-15
[338] Figure 11.35 - Front view of McDonnell Douglas F-15 Eagle twin-engine jet fighter. [mfr] [Original photo was in color, Chris Gamble, html editor]


overhead drawing of F15 wing shape
Figure 11.36 - Approximately wing-planform shape of McDonnell Douglas F-15 jet fighter.

[339] When employed as an air-superiority fighter, armament of the F-15 consists of a Vulcan 20-mm rotary cannon together with four Sidewinder and four Sparrow missiles. Although originally billed as a dedicated air-superiority fighter, the F-15 is now replacing the Convair F-106 Delta Dart as an interceptor, and trials are being made of a ground-attack version of the aircraft known as the Strike Eagle. For this latter mission, some 16 000 pounds of external ordnance can be carried. Certainly, the McDonnell Douglas F-15 Eagle will be an important part of the USAF inventory for a long time to come.
General Dynamics F-16 Fighting Falcon
In February 1972, the USAF issued a request for proposal for an experimental, lightweight, low-cost, highly maneuverable day fighter with a Mach 2.0 capability. Although there was no assurance of a follow-on production contract, five companies submitted proposals; from these, General Dynamics and Northrop were selected to build prototypes to be used in a flyoff-type competition for selection of a final winner. The single-engine General Dynamics F-16 was eventually selected over the Northrop F-17; in January 1975, the USAF announced that the aircraft would be put into quantity production. (The twin-engine Northrop F-17 later became the basis for the Navy F-18 Hornet. Descriptions of the F-18 can be found in refs. 161, 163, and 200.) Seven other countries also selected the F-16 for their use, and production lines were established in both the Netherlands and Belgium. According to reference 177, about 1800 aircraft had been built or ordered by 1980. First flight of the F-16 took place in January 1974, and it first entered squadron service with the USAF in January 1979.
The F-16 Fighting Falcon is shown in figure 11.37, and a sketch of the wing-planform shape is given in figure 11.38. The aerodynamic configuration of the F-16 is a highly integrated synthesis of such components as wing, fuselage, and inlet, with the aim of achieving maximum favorable flow interaction with subsequent optimization of overall performance. Configuration features include a cropped delta wing mounted near the top of the fuselage with large strakes extending forward from the leading edge to the sides of the fuselage. A single vertical tail is utilized together with a small fixed ventral fin located on the bottom of the fuselage (fig. 11.37). The all-moving horizontal tail is mounted in the low position and incorporates a small amount of negative dihedral.

aerial view of F-16
[340] Figure 11.37 - General Dynamics F-16 Fighting Falcon single-engine jet fighter. [mfr] [Original photo was in color, Chris Gamble, html editor]

A fixed-geometry, chin-mounted inlet supplies air to the single Pratt & Whitney F100-PW-200 turbofan engine, which is a variant of the same power plant utilized in the F-15. Since the forward portion of the fuselage provides some external flow compression, reasonable inlet efficiency is obtained even at a Mach number of 2.0. Good inlet efficiency through a wide range of angle of attack is ensured by the location of the inlet on the bottom side of the fuselage at a fore-and-aft location behind the forward intersection of the wing strakes with the side of the fuselage.
As shown by figure 11.38, the cropped delta wing blends into the fuselage sides with large strakes that extend forward from the wing leading edges. Vortexes generated by these strakes help prevent wing stall at high angles of attack and thus increase the lifting capability of the wing. Leading-edge sweepback angle is 45° and the airfoil-section thickness ratio is 4 percent. Trailing-edge flaparons serve the double purpose of high-lift flaps and ailerons for lateral control. Leading-edge maneuvering flaps are deployed automatically as a function of Mach number and angle of attack.
In some respects, the control system of the F-16 represents a complete departure from previous fighter design practice. Although conventional-type aerodynamic control surfaces are employed, the control system utilizes a novel method of transmitting pilot commands to these surfaces. In previous fighter designs, some form of mechanical device linked the control stick and the rudder pedals to the hydraulic actuating system that moved the control surfaces. In contrast, the F-16 utilizes a fly-by-wire system in which movement of the pilot's controls....

overhead drawing of F16 wing design
[341] Figure 11.38 - Approximate wing-planform shape of General Dynamics F-16 jet fighter.


...initiates electrical signals that activate the hydraulic systems and cause the control surfaces to be moved in a prescribed manner. The fly-by-wire system is lighter, simpler, and more precise than the older mechanical systems, but it does raise questions relating to electrical system reliability. In the F-16, redundancy is provided in the electrical generating and distribution equipment, and four dedicated sealed-cell batteries give transient electrical power protection for the fly-by-wire system. Two completely separate and independent hydraulic systems supply power for actuation of the aerodynamic control surfaces and other utility functions.

Another novel feature in the control system of the F-16 is the incorporation of "relaxed static stability." This means that the inherent longitudinal stability is reduced, to a level traditionally thought to be unacceptable, by moving the aircraft center of gravity to a point very near the aerodynamic center of the aircraft. Tall load and associated trim drag are reduced by this process. Compensation for the loss in inherent aerodynamic stability is provided by a combination electronic-hydraulic stability augmentation system that senses uncalled-for [342] departures from the intended flight condition and injects corrective signals into the flight control system.
Finally, the arrangement of the pilot's control stick is a radical departure from standards that trace their origin to the early days of World War I. Traditionally, the fighter pilot's control stick used for actuation of the ailerons and elevators has consisted of a lever mounted on the floor of the cockpit between the pilot's legs. (There have, of course been many variations in the detail design of the control stick.) On the F-16, the traditional control stick has been replaced by a short "side-arm controller" mounted on the right-hand console of the cockpit. The side-arm controller is a small-displacement pressure-sensitive handle that, together with the fly-by-wire system, gives the pilot the ability to exercise very precise control of the aircraft. To help prevent unwanted commands to the control handle the pilot rests his right arm in a carefully designed support. In order to increase the pilot's tolerance to forces his seat is inclined 30 in the rearward direction with
The data in table V show the design gross weight of the F-16A to be 23 357 pounds, or only about half that of the F-15C. However, wing loading and thrust- to-weigh t ratio of the two aircraft are nearly the same. Little performance information is available for the F-16A; the limited data in table V do show, however, a maximum Mach number of 2.02 at 40 000 feet and a ferry range of 2535 miles.
Originally conceived as a simple air-superiority day fighter, the aircraft was armed for that mission with a single six-barrel Vulcan 20-mm cannon and two Sidewinder missiles, one mounted at each wingtip. Over the years, however, the mission capability of the aircraft has been extended to include ground-attack and all-weather operations With full internal fuel, the aircraft can carry up to 12 000 pounds of external stores including various types of ordnance as well as fuel tanks.
The F-16 Fighting Falcon is an advanced and innovative fighter that, like the F-14 and the F-15, will be a part of the fighter scene for many years.
British Aerospace AV-8A Harrier
Discussed next is a totally unique aircraft that has an operational versatility unmatched by any other fighter in the western world. The British Aerospace Harrier can take off and land vertically like a helicopter but, unlike the well-known rotary-wing machine, accomplishes this [343] vertical-flight operation by means of a specially designed jet engine that is also able to propel the aircraft in forward flight at Mach numbers as high as 0.95 at an altitude of 1000 feet. An early prototype, known as the Hawker P-1127, flew in 1960 and was the basis of a more refined aircraft that appeared later. Known as the Kestrel, a number of these aircraft were employed during the mid-1960's in a joint military evaluation of the VTOL fighter concept conducted by the governments of the United States, the United Kingdom, and the German Federal Republic. In the 1970's, the aircraft now called the Harrier entered the active inventory of several air forces. Of the same basic design, the progression from P-1127 to Kestrel to Harrier was characterized by increased power, weight, and performance.
The Kestrel is shown in figures 11.39 and 11.40. This particular aircraft served in the joint United States, British, and German evaluation; it was later used in extensive flight studies at NASA's Langley Research Center. Today it may be seen in the National Air and Space Museum in Washington, D.C. A Harrier in service with the U.S. Marine Corps is shown in figure 11.41. The designation AV-8A is used to describe these aircraft.
The Rolls-Royce (Bristol division) Pegasus turbofan engine is the key to the great versatility of the Harrier. Unlike other jet engines with only one jet-exhaust nozzle, the Pegasus has four exhaust nozzles; two....

photo of VTOL
Figure 11. 39 - British Aerospace Kestrel single-engine VTOL jet fighter. [NASA] [Original photo was in color, Chris Gamble, html editor]


aerial view of Kestral in hover flight
[344] Figure 11.40 - British Aerospace Kestrel in hovering flight. [NASA] [Original photo was in color, Chris Gamble, html editor]


aerial photo of 2 Harrier VTOL jets
Figure 11. 41 - British Aerospace Harrier single-engine VTOL jet fighter. [ukn] [Original photo was in color, Chris Gamble, html editor]

...are located on each side of the engine. The two front nozzles discharge unheated air compressed by the fan, and the rear nozzles discharge the hot jet exhaust. A rotating cascade of vanes is used in each nozzle to vector the thrust from a horizontal direction for high-speed flight to a vertical direction for hovering and vertical takeoff and landing. Intermediate positions are used for short takeoff and landing (STOL) and [345] for maneuvering in combat situations. (This latter technique is referred to as VIFF, vectoring in forward flight.) The use of VIFF to enhance aircraft maneuverability and hence combat effectiveness was pioneered in flight studies at the Langley Research Center in the late 1960's and early 1970's. For rapid deceleration, the nozzles can actually be rotated past the vertical position to about 98°. The thrust-vectoring nozzles can be seen in the side of the fuselage in figure 11.39.
Another key element in the Harrier concept is the method for controlling the aircraft. When operated as a conventional airplane, the usual ailerons, rudder, and horizontal tail are used to generate aerodynamic control moments about the roll, yaw, and pitch axes, respectively. In hovering flight and at low forward speeds, however, the aerodynamic controls are ineffective, and reaction jets are used to provide the necessary control moments. At intermediate speeds, both reaction jets and aerodynamic controls are used. As indicated in figure 11.39, pitch jets are located at the nose and tail of the fuselage, a roll jet is at each wingtip, and a yaw jet is located behind the tail. The reaction jets utilize compressed air from the high-pressure engine compressor and respond in a proportional fashion to conventional movements of the control stick and rudder pedals. The control jets come into operation automatically when the thrust-vectoring nozzles are rotated to any angle in excess of 20°. Control of the thrust-vectoring nozzles is exercised by a lever in the cockpit located alongside the throttle.
Although the engine and reaction control system are the key elements that give unique operational capability to the Harrier, the airframe itself exhibits several interesting features. With 12° anhedral (negative dihedral), the 34° sweptback wing is mounted on top of the fuselage; like the wing, the all-moving horizontal tail has a large anhedral angle (15°). The anhedral angles of the wing and horizontal tail are intended to minimize the aircraft rolling moments due to sideslip. Even so, at certain combinations of low speed and high angle of attack, aerodynamic rolling moments greater than the combined aerodynamic and reaction control power may occur if the angle of sideslip is allowed to exceed a prescribed value. To assist the pilot in maintaining the angle of sideslip within acceptable limits, a small yaw vane that provides a visual indication of sideslip angle is mounted on the fuselage just ahead of the windshield.
The unusual landing gear of the Harrier is designed to avoid interference with the engine and thrust-vectoring nozzles. A single twowheel bogie is located in the fuselage behind the engine, and a single steerable nose-wheel is in front of the engine. Balancing outrigger [346] wheels mounted at the wingtips retract into the reaction control fairings. (See fig- 11.41.) The wing anhedral angle minimizes the length of the outrigger landing-gear struts. Also evident in the figure are the large side-mounted subsonic inlets that supply air to the 21 500-poundthrust engine.
The fighter version of the aircraft is manned by a single pilot; a two-seat trainer with the full military capability of the single seater is also available. As with so many modern jet fighters, the Harrier is equipped with zero-zero ejection seats; that is, crew escape is possible on the runway at zero altitude and zero speed.
The data in table V for the AV-8A version of the Harrier show a design gross weight of 18 000 pounds for VTOL operation and 26 000 pounds for STOL use. For the design gross weight as a VTOL aircraft, the thrust-to-weight ratio is 1.19 and the wing loading is 89.5 pounds per square foot. Maximum speed is listed as Mach 0.95 at an altitude of 1000 feet, and 2.38 minutes are required to reach 40 000 feet; service ceiling is 48 000 feet, and ferry range with maximum external fuel is 2070 miles.
Primary mission of the Harrier as employed by the Royal Air Force is that of a ground-attack fighter-bomber. In this role, a variety of external ordnance with maximum weight up to 5000 pounds may be carried, as well as two 30-mm cannons. The Royal Navy employs the aircraft in a fleet air-defense role; in this capacity, Sidewinder missiles are carried in addition to the cannon and various external stores. In naval use, the Harrier employs a short takeoff technique from a small carrier equipped with a ski-jump launching ramp; after its mission and at a much reduced weight, the aircraft makes a vertical landing on the carrier. This mode of operation is referred to as STOVL, short takeoff and vertical landing. Although generally available information is far from complete, the Harrier was apparently employed with great effectiveness in the Falkland Islands dispute between Great Britain and Argentina in 1982.
At the present time, the British Aerospace Harrier is used by the Royal Air Force and Royal Navy, the U.S. Marine Corps, and the navies of Spain and India. By mid-1980, about 304 aircraft had been produced or were on order; of this number, 110 were in service with the U.S. Marine Corps (ref. 177). An improved version of the Harrier, known as the AV-8B, is now being sought by the Marine Corps. If procured in production quantity, this aircraft will be manufactured in the United States by McDonnell Douglas under an agreement with the British Aerospace Corporation.